1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a cooled turbine rotor blade with trailing edge tip corner cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In an industrial gas turbine engine, a turbine section with a plurality of stages includes rotor blades and stator vanes to extract energy from a hot gas flow that is passed through the turbine to drive the rotor shaft. It is well known in the art of gas turbine engines that the engine efficiency can be increased by increasing the temperature of the hot gas flow entering the turbine. However, the highest temperature is limited to the material characteristics of the turbine parts, especially the first stage turbine stator vanes and rotor blades because these are directly exposed to the hot gas flow exiting the combustor.
Also in an industrial gas turbine engine (IGT), the longevity of the parts is an important design factor since these engines generally run for over 48,000 hours before shut down and inspection of parts. Any premature shutdown caused by a damaged part such as a turbine blade will result in significant increases in engine operating cost.
A first stage turbine rotor blade is exposed to the hot gas flow. A complex arrangement of internal cooling passages is used to provide cooling to the blade such that the blade can be used under extreme thermal conditions that would normally melt parts of the blade. Hot spots can occur on parts of the blade due to low levels of cooling. Hot spots can cause erosion of blade parts that will result in loss of efficiency to the engine and damaged parts that must be replaced. FIG. 1 shows a prior art conical turbine blade with a suction side cut back tip rail having a tapered tip rail tip corner.
One major problem with the prior art first stage turbine blade is in the blade trailing edge tip section. The prior art pressure side bleed tip rail design yields a suction side tip rail region which is very difficult to cool. High temperature turbine blade tip section heat load is a function of the blade tip leakage flow. A high leakage flow will induce a high heat load onto the blade tip section. thus, blade tip section sealing and cooling must be addressed as a single problem. The prior art turbine blade tip includes a squealer tip rail which extends around the perimeter of the airfoil flush with the airfoil wall to form an inner squealer pocket. The main purpose of incorporating a squealer tip in a blade design is to reduce the blade tip leakage and also to provide for rubbing capability for the blade against an outer shroud.
FIG. 2 shows a prior art blade squealer tip cooling design. Film cooling holes are formed along the airfoil pressure side tip section from the leading edge to the trailing edge of the blade to provide edge cooling for the blade pressure side squealer tip. Also, convective cooling holes are formed in the tip rail at an inner portion of the squealer pocket to provide additional cooling for the squealer tip rail. Secondary hot gas flow migration around the blade tip section is also shown by the arrows in FIG. 2. U.S. Pat. No. 5,564,902 issued to Tomita on Oct. 15, 1996 and entitled GAS TURBINE ROTOR BLADE TIP COOLING DEVICE discloses this squealer tip design.
FIG. 3 shows an enlarged view for the blade trailing edge squealer tip section of FIG. 2. Since the blade tip rail is cut off at the aft section of the pressure side (the tip rail on the pressure side does not extend all the way to the trailing edge), it becomes a single squealer tip rail configuration and thus decreases the ability to reduce the blade tip leakage flow. The suction side blade tip rail is subject to heating from three exposed sides. As a result, cooling of the suction side squealer tip rail from the row of film cooling holes along the blade pressure side peripheral and at the bottom of the squealer floor becomes insufficient. This is primarily due to the combination of squealer pocket geometry and the interaction of hot gas secondary mixing. The effectiveness induced by the pressure film cooling and tip section convective cooling holes is very limited. Therefore, for the blade trailing edge tip section, the prior art pressure side bleed tip rail design yields a suction side tip rail region which is very difficult to cool. Oxidation and erosion at the blade trailing edge suction side tip section occurs in most engine operations. Frequently this region of the blade tip needs to be rebuilt during engine overhaul cycles. This limits the life of the engine part and decreases the efficiency of the engine.
It is therefore an object of the present invention to provide for a turbine blade with a squealer tip having a trailing edge tip corner with increased cooling capability over the cited prior art references.